41results about How to "Improve flight performance" patented technology

The invention belongs to the field of trajectory optimization, and provides an unmanned aerial vehicle-unmanned vehicle joint trajectory optimization method and system, a medium and equipment, and the technical scheme is that the trajectory characteristics of an unmanned aerial vehicle and an unmanned vehicle are integrated, and the motion characteristics, environmental constraints and task requirements of the unmanned aerial vehicle and the unmanned vehicle are comprehensively considered. The cooperative task of the unmanned aerial vehicle and the unmanned vehicle is completed in a high-efficiency and high-quality manner, it is ensured that the unmanned aerial vehicle and the unmanned vehicle fully excavate the cooperative ability, unconstrained trajectory optimization is adopted, all constraint penalty terms are set as soft constraint terms, the trajectory is specifically required according to a specific task, and if the trajectory is required to pass a specific point, the unmanned aerial vehicle and the unmanned aerial vehicle are matched with each other. New constraint penalty terms can be conveniently added to the trajectory; adjustment of cooperation of an energy consumption constraint penalty term and a time regular term is added, so that the energy consumption of the unmanned aerial vehicle is reduced; landing terminal constraint is carried out on the trajectory with the landing task, so that the unmanned aerial vehicle obtains safe and efficient landing quality at the confluence point.

The utility model discloses an aircraft recovery device and an aircraft, and relates to the technical field of aircrafts. The recycling bin is installed on the foundation, the end, away from the ground, of the recycling bin is of an opening structure, and a plurality of through holes are formed in the side wall of the recycling bin in the circumferential direction; the fixing mechanisms are arranged on the outer side of the recycling bin in the circumferential direction; the multiple locking mechanisms are rotationally connected to the multiple fixing mechanisms corresponding to the locking mechanisms, at least part of the locking mechanisms penetrate through the through holes and are located in the recycling barrel, and when the aircraft is recycled, the part of the locking mechanisms located in the recycling barrel abut against the bottom of the aircraft. And the locking mechanism located outside the recycling bin penetrates through the through hole and abuts against the outer wall of the aircraft so as to lock the aircraft. Therefore, according to the aircraft recovery device, the aircraft is prevented from carrying an undercarriage, the energy consumption of the aircraft is reduced, the flight performance of the aircraft is improved, and meanwhile, the recovery safety of the unmanned aerial vehicle is improved.

The utility model provides a freight unmanned plane relates to unmanned plane technical field, including fuselage for loading goods, the fuselage tail lower portion is equipped with rear aircraft wheel, rear aircraft wheel is connected with the fuselage through support rod, rear wing is located the fuselage tail upper portion and has upper reflex angle, front wing is located the fuselage head lower portion and has lower reflex angle, the front wing both sides lower portion is equipped with front aircraft wheel, tail push propeller is located the fuselage tail, vertical propeller is located the fuselage both sides, wherein, tail push propeller drives through the fuselage upper oil dynamic piston engine, vertical propeller drives through fuselage upper motor. A freight unmanned plane in the utility model, solved the problem that the prior art lacks a freight unmanned plane capable of stable vertical take-off and landing and long range and large load.

The invention provides a jet circulation empennage control system, an empennage and an aircraft, and relates to the technical field of active flow control, and the jet circulation empennage control system comprises a circulation controller, a steering engine control module, an auxiliary power system, an empennage circulation control unit, an airflow monitoring sensor and an airflow valve. Compressed air is provided by an auxiliary power system of the aircraft, so that the weight of the aircraft can be reduced while yaw control and / or pitching control are / is realized, the number of parts of the whole aircraft is reduced, and the reliability of the system is improved. The aircraft does not need a hydraulic system with a complex structure any more, the weight of the aircraft is reduced, the consumption of fuel oil is reduced, and therefore the economical efficiency of the aircraft is improved. In addition, the take-off weight of the aircraft is reduced, the take-off and landing distance can be shortened, the effective load of the aircraft is improved, and therefore the overall flight performance of the aircraft is improved.

The present invention relates to a method for analyzing tailspin characteristics. The time history of tailspin characteristics is obtained by solving tailspin dynamics differential equations, and the tailspin characteristics of simulation examples under conditions of multi-parameter changes such as weight, moment of inertia, and entry mode are predicted, and the influence of multi-variable parameters on the tailspin characteristics is analyzed; a machine learning model is built using an XGBoost regression algorithm based on decision tree analysis, and the multi-variable parameters of the simulation example are analyzed for feature importance with tailspin characteristics such as average angle of attack, average rotation rate, and sinking speed as targets; the influence ratio of multi-variable parameters on the same tailspin characteristics is analyzed, and the influence of key factors on the tailspin is completed in a quantitative manner. The tailspin feature analysis method of the present invention takes tailspin characteristics as targets, performs feature importance analysis on three multi-variable parameters of weight, moment of inertia, and entry mode, gives the degree of influence of each variable parameter on the corresponding target in the form of quantitative weights, and completes the multi-dimensional global feature analysis of the simulation example.

This invention belongs to the field of variable wing technology and discloses a lightweight integrated aerodynamic drive system and a deformable wing. The aerodynamic drive system includes: a double-acting cylinder integrated with the valve bodies of a normally open and normally closed electromagnetic directional valve via a cylinder head; a combined valve comprising an inflation valve and a pressure reducing valve; the inflation valve is located above the pressure reducing valve, with its valve body integrated with the pressure reducing valve's valve body, and its outlet flow channel directly connected to the pressure reducing valve's inlet flow channel; the pressure reducing valve is located on the top of the front cylinder head of the double-acting cylinder, and is connected to the normally open electromagnetic directional valve via a gas-sealed connection, directly docking with a gas cylinder. This system combines optimized gas path design and hardware integration, integrating energy and drive functions, resulting in a compact and lightweight structure. It can adapt well to the requirements of limited installation space, significantly improving the system's adaptability to various operating conditions, thereby enhancing the aircraft's flight performance.

The present invention provides a control and deployment locking system for a tube-launched UAV Π-shaped folding wing, and belongs to the technical field of tube-launched UAV control. In addition to a rotation drive module for driving the main wing and the outer wings on both sides to rotate and deploy from a folded position and a deployment locking module for locking them to a fully deployed position, the system is also provided with an outer wing folding locking module for locking the outer wings on both sides to a folded position and a control module for controlling the time-sharing unlocking of the outer wings, so that the wing on one side rotating forward unlocks the outer wing when the main wing rotates to be perpendicular to the fuselage, and the wing on one side rotating backward unlocks the outer wing before the other outer wing at a predetermined time, generating a rolling restoring torque to balance the rolling torque generated by the difference in incoming flow speed on the wing surfaces on both sides at the initial stage of the main wing deployment, solving the problem that the control mode in which the outer wings are folded in the UAV launch tube by relying on the fuselage constraint and are deployed autonomously after the UAV is launched from the tube causes a large deviation in the rolling attitude of the UAV after it leaves the tube.

The application belongs to the field of insect breeding, and specifically discloses application of citric acid or capric acid in improving flight ability of irradiated sterile male codling moth, wherein the codling moth is fed with a sucrose solution containing 4 mM citric acid or a sucrose solution containing 250 mu M capric acid within 24 hours after the irradiated male codling moth emerges from the cocoon. Compared with the prior art, the application proves that the feeding treatment of citric acid or capric acid can improve the flight ability of the irradiated sterile male codling moth, and further improve the mating success rate of the sterile male codling moth and the control effect on the population.

The invention discloses a fixed time increment fault-tolerant control method for a control surface mixed fault aircraft. The method comprises the following steps: analyzing the fault type and performance of an aircraft control surface; modeling a control surface fault aircraft; improving the design of a tracking differentiator; designing an incremental dynamic inverse control law; and designing a fixed time increment dynamic inverse fault-tolerant controller. The control method provided by the embodiment of the invention is low in cost and simple in structure, can realize accurate acquisition of the angular acceleration signal and noise filtering, reduces hardware consumption of the angular acceleration sensor, and has a practical engineering application prospect; the robustness is high, and stable control and instruction tracking of the aircraft under the control surface mixed fault can be achieved; the method is high in convergence speed and good in steady-state performance, can guarantee that the faulty aircraft quickly recovers to be stable in a fixed time and accurately tracks the instruction, and effectively improves the flight performance of the faulty aircraft.

This application provides a compound variable-pitch propeller system and control method based on circulation control, relating to the field of aviation technology. The compound variable-pitch propeller system based on circulation control includes: a dual-rotor variable-pitch motor, which is disposed in the propeller hub. The dual-rotor variable-pitch motor includes a first rotor and a second rotor. The first rotor is connected to an inner worm gear, which is connected to a variable-pitch mechanism. The second rotor is connected to an outer worm gear, which is connected to a control surface deflection mechanism. The propeller blades include fixed blades and movable control surfaces. The fixed blades are connected to the propeller hub, and the movable control surfaces are rotatably connected to the fixed blades. The propeller blades have jet slits. A transmission mechanism connects the control surface deflection mechanism and the movable control surfaces for the deflection motion of the movable control surfaces, thereby improving the efficiency and adaptability of the propeller and optimizing flight performance.

The aerodynamic layout comprises a fuselage, wings and an empennage, the wings are installed on the fuselage, the empennage is connected with the wings, the wings comprise a middle wing, outer wings and winglets, the middle wing is connected to the upper end of the middle of the fuselage, the outer wings are symmetrically arranged on the two sides of the middle wing, the winglets are arranged on the edges of the outer wings, the plane shape of the middle wing is rectangular, and the plane shape of the middle wing is rectangular. The plane shape of the outer wing is trapezoid, the sweepback angles of 1 / 4 chord lines of the middle wing and the outer wing are both 0 degree, the dihedral angle of the middle wing is 0 degree, the dihedral angle of the outer wing is 0.5 degree, the empennage comprises a vertical tail and a horizontal tail, and the vertical tail is connected to the wings and arranged behind the fuselage. A stable laminar flow field can be formed in a low-speed cruising state through the combined design of specific structures of the rectangular middle wing and the trapezoidal outer wing, wingspan loads are distributed in an oval trend through the trapezoidal design of the outer wing, induced resistance can be reduced by combining the structural design of the wingtip winglet, and additional resistance caused by the sweepback angle is eliminated through the design that the sweepback angle is 0 degree.

The utility model discloses an unmanned aerial vehicle for agriculture and forestry environment monitoring, and relates to the technical field of unmanned aerial vehicles. The unmanned aerial vehicle comprises an unmanned aerial vehicle mounting base, a mounting shell is mounted at the top end of the unmanned aerial vehicle mounting base, a fixing support is mounted in the mounting shell, a fixing plate is fixedly connected to the middle of the fixing support, and a guide sliding rail is arranged at the top end of the fixing plate; the outer wall of the fixed support is slidably connected with a plurality of moving blocks, the top end of each moving block is rotatably connected with a connecting rod, the adjacent ends of the connecting rods are rotatably connected, the ends, close to the fixed plate, of the two connecting rods are slidably connected with the top end of the fixed plate through guide sliding rails, and the bottom end of each moving block is connected with a cooling fin. An electric motor is mounted on one side of the fixed bracket; according to the unmanned aerial vehicle, when the unmanned aerial vehicle is patrolled in different flight states through the adjusting assembly, the contact area between the cooling fins and air is reasonably switched, and the cooling efficiency is improved.

The present application discloses a large aspect ratio wing flutter control method and control device thereof, including: obtaining the initial flutter speed of the wing; installing a steel cable on the wing, and applying a preload force to the steel cable in flight through a servo mechanism to limit the wing torsion; obtaining the extended flutter speed of the wing after the steel cable applies the preload force; enabling the wingtip angular displacement sensor during the flight of the UAV, and the wingtip angular displacement sensor is used to monitor the wingtip torsion angle and torsion angular velocity in real time; when the flight speed of the UAV exceeds the extended flutter speed, using the servo mechanism for dynamic control. The method solves the problem of how to increase the flutter speed, so that the UAV can fly safely at a higher speed, thereby improving its flight performance.

The invention discloses a hydrodynamic rocket and a working method thereof. The hydrodynamic rocket comprises a hydrodynamic rocket body. The injection direction adjusting actuation assembly comprises a mounting cylinder connected to the tail part of the hydrodynamic rocket main body and an actuation mechanism arranged in the mounting cylinder; the actuating mechanism comprises a frame rotationally connected to the inner wall of the mounting cylinder, an injection pipe tail mounting base rotationally connected into the frame, a first driving mechanism used for driving the frame to rotate and a second driving mechanism used for driving the injection pipe tail mounting base to rotate. The rotating axis of the frame intersects with the rotating axis of the injection pipe tail mounting base in a crossed mode, and the tail of an injection pipe of the hydrodynamic rocket is connected to the injection pipe tail mounting base. The invention aims to control the flight attitude of the hydrodynamic rocket after launching.

The present application relates to a kind of micro aircraft low-speed lift increasing wrinkle airfoil, including main wing, wind speed detection sensor, shape memory alloy wire, direct current power supply, remote control machine, the main wing is the wrinkle wing profile structure with lift increasing effect, the wind speed detection sensor is installed on main wing and the wind speed signal detected is passed to remote control machine, the shape memory alloy wire is electrically connected with direct current power supply, the shape memory alloy wire and direct current power supply are installed on main wing, the direct current power supply is connected with remote control machine and is controlled by remote control machine, the on-off electricity situation of shape memory alloy wire and direct current power supply is controlled by remote control machine, make main wing switch between wrinkle wing type and flat plate wing type.The present application can significantly improve the lift of micro aircraft wing, has the advantages such as controllable deformation, reliable structure, stable driving and low energy consumption.

The invention discloses an artificial feather leaf for a badminton, relates to the field of manufacturing of sports goods, and particularly relates to a specially designed artificial feather leaf structure which is suitable for the badminton. The invention provides an artificial feather leaf for a shuttlecock, which is characterized in that the artificial feather leaf consists of a feather piece and a feather vein; wherein the feather vein is used as a skeleton, and the feather piece and the feather vein form an artificial feather leaf. The artificial feather leaf is composed of an upper layer, a middle layer and a lower layer, the upper layer and the lower layer are feathers which are bonded through glue, and the feather vein serves as a framework and is clamped between the two layers of feathers. The feather veins are preferably made of titanium-nickel memory alloy with hyperelastic performance, and the feather pieces are made of plastic films. The super-elastic titanium-nickel memory alloy is adopted as the framework, the deformation capacity of the artificial feathers at different flight speeds is remarkably enhanced, the ball hitting hand feeling is further optimized, and meanwhile the flight characteristics of the artificial feathers are closer to those of natural badminton balls.

The invention discloses an aircraft linear driving device which comprises a shell, a motor and a lead screw are arranged in the shell, the output end of the motor is connected with the lead screw through a connecting piece, a lead screw nut is installed on the outer side of the lead screw, a sleeve is installed on the lead screw nut, a top head is detachably installed at one end of the sleeve, and the top head is detachably installed on the other end of the sleeve. A plug matched with the sleeve is installed at one end of the shell. With the structural characteristic of vertical extension and retraction of the push rod, the device has the characteristics of attractive appearance, small size, low transmission resistance, low noise, convenience in installation and the like; the aircraft has the precise control capability, high-precision power output can be provided, precise control over all parts of the aircraft, including a control surface, a flap, an undercarriage and the like, is achieved, and therefore the maneuverability and the flight performance of the aircraft are improved.

The utility model belongs to the technical field of unmanned plane, concretely relates to an unmanned plane brake wheel, including fork arm, wheel shaft, wheel hub, tyre, bottom cover, brake drum, brake shoe and rocker, the utility model discloses a brake component is integrated in the wheel cavity of wheel hub, has reduced the volume and weight of overall structure, meets the requirement of unmanned plane to light weight and compact structure, has improved the flight performance and endurance of unmanned plane, again adopts the brake mode of brake shoe and brake drum friction, can provide greater braking force, ensures that unmanned plane can stop rapidly and effectively when needing to brake, and realizes the transmission of brake line tension to brake shoe action through rocker, rotating rod and other components, and its structure is simple, transmission efficiency is high, reduces the possibility of energy loss and fault occurrence, and is easy to maintain and overhaul, reduces use cost.

The embodiment of the utility model provides a battery device, an energy storage device and an electric device. The battery device comprises a battery monomer and a phase change piece, the battery cell includes a pressure relief mechanism. The multiple phase change pieces are stacked in multiple layers in the first direction to form a stacked body, the stacked body and the pressure relief mechanism are oppositely arranged in the first direction, each phase change piece comprises a shell and a phase change material, a containing cavity is formed in the shell, the phase change materials are stored in the containing cavity, and the shell can be broken under the heat effect of emissions discharged by the pressure relief mechanism to release the phase change materials; and the other peripheral phase change pieces are displaced towards the position of the cracked phase change piece due to the loss of the support of the cracked shell. According to the battery device provided by the embodiment of the invention, after part of the phase change pieces in the stack body are broken under the heat action of the emissions sprayed out by the pressure relief mechanism, other phase change pieces can generate displacement to fill gaps generated by the broken phase change pieces, so that the temperature of the emissions can be better reduced.

The invention relates to an ultrahigh-pressure-ratio compact efficient energy-saving turbojet engine which comprises a main body part, a two-stage supercharging unit, a three-stage supercharging combustion unit and a reel structure connecting pipeline system. The reel structure connecting pipeline system comprises a first reel structure connecting pipeline, a second reel structure connecting pipeline, a third reel structure connecting pipeline and a fourth reel structure connecting pipeline. Along with the increase of the supercharge ratio, combusted gas expands and does work more sufficiently in the turbine, heat energy generated after fuel is combusted is converted into mechanical energy more effectively, the fuel oil heat efficiency is greatly improved, the fuel oil economical efficiency is improved, and fuel oil consumption is greatly reduced.

The application relates to a variable sweep telescopic wing structure, which comprises a wing platform and a wing assembly, the wing assembly comprising a first wing and a second wing, a first driving assembly, a rotating assembly and a lifting shaft, the rotating assembly comprising a first rotating shaft and a second rotating shaft which are connected in parallel on the wing platform, the first rotating shaft being coaxially sleeved on the lifting shaft, the first rotating shaft being fixedly matched with the lifting shaft in the circumferential direction and being slidably matched with the lifting shaft in the axial direction, the first wing being connected at the top of the lifting shaft, the second wing being connected at the top of the second rotating shaft, and the first wing and the second wing both being telescopic structures, the first driving assembly being transmissionally connected with the first rotating shaft and the second rotating shaft to drive the first rotating shaft and the second rotating shaft to rotate and drive the lifting shaft to lift. The structure can realize wing unfolding and folding with any sweep angle, increase the stability and flexibility of the aircraft, be telescopic to change the wing area, adapt to different flight states, improve the wing efficiency and improve the flight performance.

The invention relates to the technical field of aviation aircraft design, and discloses a variable camber wing based on a bistable composite material structure driving multi-section rotation, which comprises a wing leading edge structure, a multi-section rotation rigid skin, an internal truss structure connected with the multi-section rotation rigid skin, a bistable composite material driving structure and a transmission mechanism, the bistable composite material driving structure is arranged between the rigid skin hinge and the transmission mechanism; and the transmission mechanism is arranged between the truss structure in the rigid skin and the sliding chute of the wing leading edge structure. The control precision and reliability are improved through relative change of the three rigid wings, and the problems that in the prior art, the structure is complex, the weight is large, the actuation efficiency is low, and the bearing capacity is low are solved.

The invention provides an aircraft controlled by combination of flexible wings and grid rudders. The aircraft comprises an aircraft body, the grid rudders and the flexible wings, the grid rudder is arranged at the tail end of the aircraft main body, and attitude control is performed on the aircraft through deflection of the grid rudder; the flexible wings are installed on the side wall of the aircraft body and located on the sides, away from the tail end of the aircraft body, of the grid rudders, and the flexible wings are unfolded and changed within the unfolding angle range. Through the installation of the flexible wings, the problems that the control capability of the grid rudder is insufficient in the transsonic speed stage and the pressure center coefficient is obviously changed are effectively solved, and the static stability of the aircraft is improved by controlling the change of the unfolding angle of the flexible wings. According to the aircraft, efficient and stable control can be achieved in the full-speed domain, compared with a traditional aircraft, the aircraft has the advantage of being lower in resistance during subsonic-speed and supersonic-speed flight, and good stability and high maneuverability are shown in the transonic-speed stage.

The invention provides a multi-rotor bomb-dropping fire extinguishing unmanned aerial vehicle paddle optimization design method, and relates to the technical field of unmanned aerial vehicles, the method comprises the following steps: obtaining fuselage geometric design data and paddle design data, using CFD software to construct a fire extinguishing unmanned aerial vehicle model, using the model to carry out aerodynamic analysis, and obtaining flight power data; generating a blade flight evaluation coefficient according to the blade material property data and the flight power data; obtaining a blade flight evaluation coefficient under each rotating speed gear, and analyzing the blade flight evaluation coefficient to obtain a flight quality evaluation coefficient; constructing and training a flight quality evaluation model by taking the historical blade design data as input and the corresponding historical flight quality evaluation coefficient as a label; and by taking the blade design data as an initial solution and taking blade flight evaluation coefficient maximization as an optimization target, optimizing the blade design data by using a simulated annealing algorithm to obtain optimal blade design data.

The present invention relates to the technical field of scaled verification aircraft, specifically to a scaled verification aircraft with an active double-throat jet vector nozzle, comprising: an airframe structure, wherein a flight combination control system is arranged inside the airframe structure; a landing mechanism; an emergency parachute recovery mechanism, comprising a parachute recovery unit, wherein a steering gear controller of the parachute recovery unit is connected to the flight combination control system in communication; a servo rudder system, which is connected to the flight combination control system in communication; a digital navigation system, which is used to exchange flight data of the verification aircraft during operation and control the navigation information of the verification aircraft in real time with the flight combination control system; a power system, comprising a power mechanism, an oil supply mechanism, a jet mechanism and a vector nozzle structure, wherein the power mechanism is connected to the flight combination control system in communication, and the vector nozzle structure adopts an active double-throat jet vector nozzle structure. The present invention solves the problem that the existing verification aircraft is usually verified in a wind tunnel environment and cannot obtain the actual flight environment.

The present invention proposes a method and system for calculating the lateral deviation of a landing runway, and relates to the field of civil aviation safety technology. The method includes: obtaining the corresponding relationship between the runway distance and the offset angle and the localizer deviation according to the landing runway design specification; the runway distance is the distance from the runway entrance to the localizer beacon; obtaining the runway entrance marking line, and reversely looping the flight motion trajectory according to the flight QAR data and the flight landing point; obtaining the runway entrance height based on the intersection of the flight motion trajectory and the runway entrance marking line; obtaining the localizer deviation parameter corresponding to the runway entrance height, and calculating the lateral offset distance based on the runway distance and the corresponding relationship. By identifying the position where the aircraft reaches the runway entrance based on the reverse loop of the landing point, the accurate lateral deviation is solved according to the localizer deviation and runway distance at the entrance position, thereby improving the accuracy of the calculation result; at the same time, the parameters involved are only based on the landing runway design specification, and there is no need to associate the aircraft model information, so the versatility is stronger.

The application discloses a kind of unmanned plane recycling methods, comprising the following steps: ground station system obtains the position information of fixed-wing unmanned plane, keeps the flight state of fixed-wing unmanned plane, and issues take-off instruction to multi-rotor unmanned plane;Multi-rotor unmanned plane tracks the navigation point of fixed-wing unmanned plane, shortens the relative distance between multi-rotor unmanned plane and fixed-wing unmanned plane;Adjust the flight parameters of multi-rotor unmanned plane, keep the distance difference of multi-rotor unmanned plane and fixed-wing unmanned plane in height direction;Adjust the relative motion distance between multi-rotor unmanned plane and fixed-wing unmanned plane, and carry out docking recycling;Close the power system of fixed-wing unmanned plane, control multi-rotor unmanned plane to land to preset place.The application can well utilize multi-rotor unmanned plane to realize the recycling of fixed-wing unmanned plane in the air.The application also discloses a kind of unmanned plane recycling system, ground station system, unmanned plane.

The invention relates to the technical field of power transmission, in particular to an unmanned aerial vehicle remote power transmission system which comprises an unmanned aerial vehicle body, sending end equipment and receiving end equipment, strong current is generated through the sending end equipment, a magnetic field or an electric field is generated through the strong current, and the magnetic field or the electric field is received through the receiving end equipment and converted into power. The unmanned aerial vehicle provides power for the unmanned aerial vehicle body, so that wireless transmission of the power can be realized, the voyage of the unmanned aerial vehicle can be increased, the continuous operation time of the unmanned aerial vehicle in the air can be prolonged, the unmanned aerial vehicle can fly farther, the flight radius or the task execution range can be expanded, and meanwhile, the flight capability and maneuverability of the unmanned aerial vehicle can be enhanced; according to the utility model, the wireless power transmission system is simple in structure, reduces the dependence on the ground base, reduces the logistics support work of the ground base, and solves the problems that the mobility and flexibility of equipment are limited because most of the existing power transmission modes depend on wired connection, and energy loss and potential safety hazards are caused by the problems of line aging, poor contact and the like.